This disclosure relates to gas turbine engines, and more particularly to film cooling of turbine components of gas turbine engines.
Advancements in performance of turbomachines, such as gas turbine engines, performance has often been linked to overall pressure ratio of the turbomachine and a turbine inlet temperature that can be sustained during operation of the turbomachine. Increases in efficiency through increases in pressure ratio and/or turbine inlet temperature typically results in an increase in operating temperatures of turbine flow path components, in which temperatures of the working fluid in the turbine flow path is often several hundred degrees Fahrenheit higher than the melting point of component materials.
Cooling air is often extracted from lower temperature portions of the turbomachine, for example, the compressor, and is utilized to cool the turbine components. One type of cooling utilized with this cooling flow is film cooling where the cooling air is delivered to an interior of the component then emitted over an external surface of the component. The cooling air is typically emitted through cooling holes that are machined into the part, and are circular in cross-section. Diffusion shapes are often added around the hole at the external surface.
Diverting the cooling air from the compressor incurs efficiency penalties that increase with the increase in cooling air use due to increases in pressure ratio and/or turbine inlet temperature. Thus, to reduce the efficiency penalty, it is desired to reduce the necessary cooling flow by increasing film cooling effectiveness.